Supersonic combustion engine

ABSTRACT

1. An airbreathing turbine-type powerplant comprising: A SUPERSONIC CONVERGENT INLET HAVING A THROAT IN WHICH AIR FLOWS AT SUPERSONIC VELOCITY, A PLURALITY OF COUNTERROTATING SUPERSONIC IMPULSE ROTORS LOCATED AT THE THROAT OF SAID INLET WHEREBY THE AIR IS ACCELERATED AND DISCHARGED AT SUPERSONIC VELOCITY MEANS FOR DRIVING SAID ROTORS, A PLURALITY OF SUPERSONIC STRAIGHTENING STATORS DOWNSTREAM OF SAID ROTORS, MEANS FOR INJECTING FUEL INTO SAID SUPERSONIC AIRSTREAM THROUGH SAID SUPERSONIC STRAIGHTENING STATORS, A FIXED, ANNULAR SUPERSONIC DIFFUSER DOWNSTREAM OF SAID STATORS, AND A COMBUSTION CHAMBER DOWNSTREAM OF SAID FUEL INJECTING MEANS FOR COMBUSTING SAID FUEL-AIR MIXTURE AT SUPERSONIC VELOCITIES.

United States Patent Hausmann et al.

[ Mar. 19, 1974 1 SUPERSONIC COMBUSTION ENGINE [75] Inventors: George F.Hausmann, Glastonbury,

Conn.; Arthur W. Blackman, Cambridge, Mass.

[73] Assignee: United Aircraft Corporation, East Hartford, Conn.

[22] Filed: Oct. 24, 1965 [2]] Appl. No.: 504,706

52 US. Cl 60/268, 60/224, 60/246, 60/39.43

51 Int. Cl. F02k 3/00 58 Field of Search 60/224, 246, 268, 269,

[56] References Cited UNITED STATES PATENTS 2,509,890 5/1950 Stacker60/224 2,579,049 12/1951 Price 60/39.35

3,111,005 11/1963 Howell et a1 60/246 3,203,180 8/1965 Price 60/2623,237,400 3/1966 Kuhrt 60/246 Primary Examiner-Samuel Feinberg Attorney,Agent, or Firm-Donald F. Bradley EXEMPLARY CLAIM 1. An airbreathingturbine-type powerplant comprising:

a supersonic convergent inlet having a throat in which air flows atsupersonic velocity,

a plurality of counterrotating supersonic impulse rotors located at thethroat of said inlet whereby the air is accelerated and discharged atsupersonic velocity means for driving said rotors,

1 Claim, 8 Drawing Figures PATENTED IIAR I 9 I974 SHEET 1 OF 2 FIG. 3

DISTANCE TNVENTORS BY Abmm a.

ATTORNEY 1 SUPERSONIC COMBUSTION ENGINE This invention relates toturbine type engines, and specifically to improved lightweight turbineengines designed to accelerate through subsonic and supersonic flightspeeds and cruise at supersonic speeds. More particularly the inventionencompasses engines having a supersonic inlet, a supersonic impulsecompressor located at the inlet throat, and a combustor operable at thesupersonic Mach numbers provided by the impulse compressor.

At the present advanced state of the art of turbine engine components itis expected that only modest gains can be realized in componentefficiencies, resulting in only slight improvement in aircraft range andpayload for such higher component efficiencies. However, reductions inpropulsion system weight will produce increases in aircraft range and/orpayload by direct substitution of propulsion system weight savings forextra payload and/or fuel. This aspect of system design is particularlysignificant since at present engine weight is nearly equal to aircraftpayload for supersonic aircraft. Further, present air induction systemsare substantially equal to engine weight for flight in the vicinity ofMach 3, and as flight speeds increase the weight and complexity ofefficient inlets increases. It is therefore obvious that large gains inthe payload or range of supersonic aircraft will result if thepropulsion system weight can be reduced by integrating the inlet,compressor and combustor components.

The use of supersonic compressors in turbo-type propulsion systems forflight at supersonic speeds offers the potential advantages of lightweight because of the high work output per stage and design flexibilityof the supersonic induction system which results from the ability of thesupersonic compressor to accommodate supersonic axial Mach numbers. Theprimary disadvantage of conventional supersonic compressor systems ofeither the shock-in-rotor or shock-in-stator types is the relatively lowefficiency which results from shockboundary layer interactions,nonequilibrium radial flows, and secondary flows within the blading.These loss-producing mechanisms are all attributed to severe adversepressure gradients within the blading.

As taught in U.S. Pat. No. 2,947,139, entitled By- Pass Turbojet, issuedAug. 2, 1960, to George F. Hausmann and assigned to United AircraftCorporation, the details of the patent teaching being herebyincorporated as part of this application, the supersonic impulse rotorimparts angular momentum and increased kinetic energy to the airflowwith little or no increase in static pressure. The airflow is thenredirected to the axial direction by either counterrotating rotorconfigurations or by impulse stators having little or no static pressurerise within the passages. For a conventional supersonic through-flowengine with subsonic combustion, an annular diffuser configuration isused to convert the resulting kinetic energy to static pressure and toreduce the flow Mach number to subsonic velocities. The diffuser may beprovided with bleed passages or a variable geometry throat toaccommodate matching requirements when applied to turbotype propulsionsystems. By utilizing the supersonic combustion teachings of thisinvention, the diffuser downstream of the compressor may be eliminatedsince there is no need to lower the Mach number below supersonic beforethe airflow from the compressor enters the combustion chamber.

The turbine engines described herein reduce the weight of the propulsionsystem by utilizing a supersonic inlet which only partially compressesthe free stream supersonic flow, thus eliminating the need for a complexinlet forfall flight regimes. A supersonic impulse compressor is locatedat the inlet throat and is designed to operate with a supersonic axialMach number. Combustion downstream of the impulse compressor is effectedsupersonically, thus eliminating the need for a complete supersonicdiffuser or stator to reduce the flow Mach number to subsonic values asrequired for conventional engines having subsonic Mach numbers.

It is therefore an object of this invention to provide a novel turbojetengine utilizing a coupled supersonic inlet, supersonic impulsecompressor, supersonic combustor and supersonic turbine.

Another object of this invention is a propulsion system for supersonicspeeds and high altitudes having lower weight than conventionalpropulsion systems.

A further object of this invention is a turborocket in which combustiontakes place at supersonic velocities.

Another object of this invention is a turbine engine in which inletweight and complexity are considerably reduced.

A still further object of this invention is a turborocket engine systemin which the turbine is driven by exhaust products from a gas generatorseparate from the combusted air. I

Another object of this invention is a novel airliquefying turborocketutilizing supersonic compression and combustion. I

A further object of this invention is a novel bypass supersoniccombustion engine.

These and other objects and a fuller understanding of this invention maybe had by referring to the following description and claims, read inconjunction with the ac companying drawings, in which:

FIG. 1 shows a conventional supersonic through-flow compressor subsoniccombustion engine; and

FIG. 2 shows a supersonic through-flow compressor turbojet engine withsupersonic combustion and a supersonic turbine; and

FIG. 3 shows a novel lightweight turborocket engine utilizing asupersonic impulse compressor and supersonic combustion together withMach number variations through the turborocket; and

FIG. 4 shows a preferred supersonic compressor blade configuration; and

FIG. 5 shows typical vector diagrams of flow velocities through thesupersonic through-flow compressor; and

FIG. 6 shows an alternate turborocket configuration; and

FIG. 7 shows a bypass engine configuration utilizing the supersonicimpulse compressor, supersonic combustion concept; and

FIG. 8 shows an air-liquefying air-turborocket.

A conventional supersonic turbojet engine employing a supersonicthrough-flow impulse compressor and subsonic combustion is shown inFIG. 1. Referring to the FIGURE, counterrotating supersonic impulserotors 10 and 12 compress the airflow from diffuser 14 whereby the flowleaves the compressor supersonically with neglibible exit swirl. Asupersonic straightening compressor stator 16 may be incorporated toreturn the flow from the compressor to the axial direction, depending onwhether or not flow with tangential velocities in the combustor isdesired.

The rotors and 12 are driven conventionally by shafts which are in turndriven by turbine rotors l8 and 20. The second impulse rotor 12 isdriven by the first stage turbine 18 and the first impulse rotor 10 isdriven by the second stage turbine 20. The use of supersoniccounter-rotating impulse rotors provides an efficient means foraccelerating the air without necessarily increasing the static pressureat this stage of the powerplant cycle.

Downstream from the compressor is an axisymmetric supersonic diffuser 22for converting the flow kinetic energy to static pressure prior to entryinto the burner 24. Bleed passages 26 having bleed doors or theequivalent are provided to permit the annular diffuser 22 to swallow theflow during starting and offdesign operation.

The location of the compressor at the inlet capture station permits anappreciable reduction in overall engine length and complexity.

Although the configuration of FIG. 1 shows a counterrotating compressorwith two shafts, a lighter design with somewhat reduced performance maybe provided by the use of a single shaft with impulse stators located atthe exit of each rotor.

If a supersonic through-flow turbojet and a conventional turbojet havinga subsonic compressor are both designed to supply the same Mach 3thrust, the frontal area of the supersonic through-flow engine would beapproximately equal to that of the turbojet and each engine would passthe same airflow. However, transonically the supersonic through-flowengine passes 35 percent more relative weight flow than the conventionalturbojet because of the larger flow area of the throughflow compressor.Thus during transonic acceleration the airflow of the supersonicthrough-flow engine would be approximately 35 percent greater than aconventional turbojet having the same Mach 3 thrust and approximatelythe same inlet diameter. As a consequence it is not necessary to employafterburning to attain adequate takeoff or transonic accelerationthrust.

A modification of the supersonic through-flow compressor engineutilizing supersonic combustion is shown in FIG. 2. The comments notedabove apply to this configuration also except as indicated. Thesupersonic combustion through-flow engine consists of counterrotatingimpulse compressor stages 28, 30 from which the flow leavessupersonically with negligible exit swirl, and which are driven bycounterrotating turbine stages 32, 34. A supersonic straightening stator36 may be incorporated. Fuel is injected through a multiplicity ofnozzles 38 or through the straightening stators 36 into the exit flowfrom thecompressor. Ignition is achieved by conventional piloting. Theturbine blades are cooled by conventional film or transpiration cooling.Downstream from the compressor is an annular supersonic diffuser 40which partially converts the flow kinetic energy to pressure and reducesthe supersonic Mach number entering the burner. The diffuser 40 isshortened relative to the diffuser in a subsonic combustion engine, andthe supersonic combustion in combination with a supersonic turbineresults in a considerably shortened engine configuration. Flow may passfrom the turbine into an afterburner (not shown) where additional fuelis injected through nozzles and the flow then exhausts through anexhaust nozzle. Bleed doors 42 or the equivalent may be provided topermit the annular diffuser 40 to swallow flow during starting andoff-design operation.

To achieve efficient supersonic combustion, sufficient combustionchamber length must be provided to meet two basic requirements. First,there must be sufficient length to provide for the ignition delay timeof the fuel-oxidizer mixture; second, the length must be compatible withthe achieving of approximately flat transverse temperaturedistributions. The first length requirement is fixed by chemistryconsiderations, and the second requirement is fixed by fluid mechanicalconsiderations. For a given fuel-oxidizer mixture and flight condition,the ignition length is an exponentially dependent function of the statictemperature of the flow stream. Hence, the ignition distance can be keptshort by providing piloting, low blockage flameholders, or upstreaminjection which raises the local static temperature. The mixing lengthcan be kept short by providing a large number of transverse ignitionsources, e.g., the mixing distance for two sources is approximately halfthe distance required for one source.

The best fuels for use with supersonic combustion engines are thosewhich have the shortest ignition delay times, the highest density, andhighest theoretical specific impulse. The choice of an optimum fuel fora specific application will depend upon the trade-offs involved sinceall fuels do not exhibit all of the desired characteristics. In mostcases either hydrogen or hydrocarbon fuels will probably be preferred.

The requirements noted above relative to the combustion chamber and fuelare pertinent to the various supersonic engine configurations describedherein. Background material relating to the theory of supersoniccombustion in ramjets, and considerations of combustion and mixingrequirements may be found by referring to articles in Astronautics andAeronautics magazine entitled Hypersonic Airbreathing Propulsion by W.H. Avery and G. L. Dugger, June l964 at page 42, and SupersonicCombustion Progress by A. Ferri, August 1964 at page 32.

Advantages of a supersonic impulse compressor, supersonic combustion andsupersonic turbine engine are: lightweight, mechanical simplicity, shortlength, a reduced number of stages, high flow handling capacity duringtakeoff and transonic acceleration, minimized dissociation of theexhaust products in the combustion chamber, and reduced pressure andheat transfer in the combustion chamber.

FIG. 3 shows a novel light weight turborocket engine system utilizing asupersonic impulse compressor and supersonic combustion which isdesigned for cruise at supersonic speeds and extremely high altitudes.FIG. 4 shows a preferred compressor blade configuration, and FIG. 5 is avector diagram showing flow velocities through the compressor blades. Arepresentative illustration of flow Mach number variation through theengine is also shown in FIG. 3.

The primary features of this engine are a light weight supersonicimpulse compressor located at the inlet throat, and a combustor operableat the supersonic Mach numbers provided by the impulse compressor. Asshown in the FIGS. 3 and 4, the compressor consists of a plurality ofhighly curved blades 50. The blades preferably have a close spacing andan annulus contraction so designed that only kinetic energy is added tothe fluid with high efficiency. The compressor has a series ofsupersonic straightening stators 52 to return the flow from thecompressor to the axial direction. The compressor is located in thethroat region of a precompression supersonic inlet 58 (FIG. 3) such thatthe compressor operates with a supersonic axial inlet Mach number.

Fuel to the supersonic combustor may be injected through support struts56 upstream of the compressor, through injectors 60 mounted on thestraightening stators 52 downstream from the compressor, or byconventional fuel injectors or nozzles in the combustor. A highlyreactive fuel such as hydrogen either alone or in combination with otherfuels and/or oxidizers, either cryogenic or storable, or a liquidslurry, would be used with this propulsion system. Combustion takesplace with supersonic velocities, and the Mach number of the flow withinthe combustor decreases as the temperature is increased.

Referring to FIG. 3, air enters diffuser 58 and passes into thecompressor 50. The air is compressed supersonically and leaves thecompressor at a Mach number greater than unity as shown in the graphconnected with FIG. 3. Fuel is injected into the airstream throughinjectors 60 on the trailing edge of the compressor stator 52 as seenalso in FIG. 4. The Mach number of the airstream is then increased in asecond diffuser passage 54. The flow is then expanded through a nozzle64.

The compressor 50 is driven by turbine 66 through a gearbox 68. Turbine66 is driven by the exhaust products from a gas generator 70 which arepassed through nozzle 72. The gas is further expanded through nozzle 74.

Fuel-air mixtures which have low ignition delays will perform best inthis engine. A possible combination is the use ofa hydrogen-low mixturein the gas generator 70 and hydrogen-air mixtures in the outer diffuser54. It is necessary to provide sufficient mixing length downstream ofthe supersonic compressor so that the residence time of the fuel-airmixture in the mixing section is equal to or greater-than the ignitiondelay of the mixture;

FIG. 5 illustrates diagrammatically the flow through the single stagesupersonic impulse rotor 50 and stator 52 by means of vectors whereinthe subscripts refer to the reference numerals adjacent FIG. 4. At thedesign condition, the inlet axial velocity is shown by vector V,,, whichmay be subsonic or supersonic. The rotor speed, U, is combined with theinlet velocity V which is essentially tangent to the upper surface ofthe rotor blade leading edge. Within the rotor 50 the flow is turned,without change in pressure or velocity, to an angle which provides anabsolute exit velocity V when the rotor speed U is subtracted from therelative blade exit velocity V In this process the velocity is increasedto a high supersonic value with no change in static pressure. Stator 52straightens the flow from the rotor to the axial direction.

The graph under FIG. 3 shows the Mach number variation within the engineof FIG. 3 when a supersonic diffusion flame is used.

FIG. 6 illustrates an alternative turborocket configuration. The generalfeatures of the engine are similar to the engine of FIG. 3. Referring toFIG. 6, the power for turbine 80 is supplied by the expansion ofregeneratively heated high pressure fuel or fuel-rich mixtures such aspure hydrogen in gas generator 82. Heating of the fuel may beaccomplished by passing the fuel from the fuel supply (not shown)through the wall of the engine adjacant the exhaust 83 before expandingthe gas past turbine 80. Compressor 84 is driven by the turbine throughgearbox 86.

After passing through the turbine 80, the fuel is fed through passage 88and injected into the airstream through injectors (not shown) mounted onstruts 90 located at the compressor inlet. Air enters diffuser 92 andpasses through the supersonic impulse compressor 84, is converted topressure in diffuser 94 and the fuelair mixture is supersonicallycombusted by conventional combustors (not shown) at 95. Stator 96 may beincluded as described previously.

FIG. 7 illustrates the application of the supersonic compressor,supersonic combustion concept to a bypass engine configuration. The mainengine assembly is similar to a conventional turbojet bypass engineconfiguration. Air is injected through inlet nozzle 100 and diffuser 102and passes through a number of stages of su personic impulse rotors 104.After passing through supersonic diffuser 106, the air is supersonicallyor subsonically combusted in burner 108. The combustion products driveturbine stages 110, which in turn drive compressor rotors 104. Anexhaust nozzle is provided as shown and afterburning may be performed.

The novel feature in this configuration is in the bypass section 112 ofthe engine. The trip of compressor blade 114 extends into the bypassportion of the engine and is operated at supersonic Mach numbers. Fuelis injected through stator 116. The air entering diffuser 118 iscompressed supersonically by rotor 114 and the fuel-air mixture isburned supersonically in the bypass airstream by conventional means (notshown) as discussed previously.

FIG. 8 shows an air-liquefier turbojet utilizing supersonic combustion.Air entering engine inlet 120 is divided into two portions. The mainportion of the airstream, approximately percent of the air, is passedthrough a single-stage supersonic impulse compressor 122 and impulsestator 124. The bleed portion of the air stream, approximately 20 of theair, enters passage 126 in inlet where the air is passed through a heatexchanger 128. A source of liquid hydrogen 130 supplies the coolantthrough line 132 to the heat exchanger 128, and the air passing throughthe heat exchanger 128 is liquefied. The liquefied. air is then pumpedthrough line 134 into a rocket chamber 136, and the hydrogen from theheat exchanger 128. is also pumped to rocket chamber 136 through line138. The liquid air and the liquid hydrogen react in rocket chamber 136,and the hydrogen-rich reaction products expand through a nozzle 140 todrive turbine 142. The reaction takes place at a relatively highpressure, 100 to 300 psi, and the reaction products include considerableI-I some N and H 0.

After being expanded through turbine 142, the turbine being on the samerotor blade as compressor 122, the hydrogen-rich reaction products arepassed through line 144 and injected supersonically into the supersonicmain airstream downstream of compressor 122 through stator 124 whichalso acts as a fuel injector. Supersonic combustion by conventionalmeans then occurs, and the combustion products are exhausted through atruncated exhaust nozzle 146.

Since an extremely low temperature coolant is required to liquefy theairstream, hydrocarbon fuels cannot be utilized with this engine, andliquid hydrogen is the preferred fuel.

It is obvious that a second stage counterrotating combinedcompressor-turbine rotor may be used with this engine as described inconjunction with previous embodiments. Likewise the turbine could be aseparate rotor or series of rotor stages driven by the expansionproducts of the liquefied air-liquid hydrogen combustion, with thecompressor being driven by the turbine and the hydrogen-rich turbineexhaust products injected downstream of the compressor. In either casethe compressor could include two counterroating stages. Further, thepumps required for the liquid hydrogen and/or liquid air could be drivenfrom the turbine-compressor shaft.

Although this invention has been described in its preferred embodiment,it will be obvious to those skilled in the art that various changes andmodifications can be made in the structure and operation of thisinvention said stators, and a combustion chamber downstream of said fuelinjecting means for combusting said fuel-air mixture at supersonicvelocities.

1. An airbreathing turbine-type powerplant comprising: a supersonicconvergent inlet having a throat in which air flows at supersonicvelocity, a plurality of counterrotating supersonic impulse rotorslocated at the throat of said inlet whereby the air is accelerated anddischarged at supersonic velocity means for driving said rotors, aplurality of supersonic straightening stators downstream of said rotors,means for injecting fuel into said supersonic airstream through saidsupersonic straightening stators, a fixed, annular supersonic diffuserdownstream of said stators, and a combustion chamber downstream of saidfuel injecting means for combusting said fuel-air mixture at supersonicvelocities.